Gas turbine engine

ABSTRACT

A gas turbine engine includes an engine core having high and low pressure compressors and high and low pressure turbines. The engine further includes a fan coupled to a low pressure shaft and the low pressure turbine by a reduction gearbox. The low pressure compressor includes no more than two compressor stages, and the low and high pressure compressor together define a cruise overall core pressure ratio of between 30 and 50.

The present disclosure relates to a gas turbine engine.

According to a first aspect there is provided a gas turbine enginecomprising a gas turbine engine core comprising high and low pressurecompressors and high and low pressure turbines, the high pressurecompressor and high pressure turbine being coupled by a high pressureshaft, and the low pressure compressor and low pressure turbine beingcoupled by a low pressure shaft, the engine further comprising a fancoupled to the low pressure shaft by a reduction gearbox, wherein thelow pressure compressor comprises no more than two compressor stages,and the low and high pressure compressor together define a cruiseoverall pressure ratio of between 30 and 50.

It has been found by the inventors that this combination of featuresprovides a highly efficient gas turbine engine, which has a relativelyshort length, and so low mass, compared to alternative configurations.

The high pressure compressor may define a cruise overall pressure ratioof between 16:1 and 27:1, and may be define a cruise overall pressureratio of between 17:1 and 20:1. Consequently, the high pressurecompressor provides for a majority of the pressure rise in the core.This contributes to the overall short length of the engine.

The high pressure compressor may have no fewer than 8 stages and no morethan 12 stages, and may consist of 9 or 10 stages. The results in anengine core having a relatively high overall pressure ratio, and arelatively low stage loading, which in turn results in highthermodynamic efficiency.

The low pressure compressor may comprise a cruise pressure ratio ofbetween approximately 1.5:1 and 2:1.

The engine may comprise a core inlet passage extending between a coreinlet and an inlet of the low pressure compressor.

A first radius change ΔR₁ may be defined by a difference between aradius R_(inlet) of the core inlet passage measured at a mid-height ofthe leading edge of an engine section stator aerofoil at an axiallyforward end of the core inlet passage, and an outlet radius R_(outlet)of the core inlet passage measured at a mid-height at a leading edge ofa first rotor stage of the low pressure compressor. A first duct loadingΔR₁/L₁ may be defined by a ratio of the first radius change ΔR₁ to afirst axial length L of the inlet duct between the axially forward endand the leading edge of the first stage of the low pressure compressor.The first duct loading may be between 0.3 and 0.6, may be between 0.35and 0.55, and may be approximately 0.4.

A first compressor rotor of the low pressure compressor may define a hubto tip ratio defined by a radius of a tip of a leading edge of thecompressor rotor divided by a radius of a root at the leading edge ofthe compressor. The hub to tip ratio may be between 0.6 and 0.75, andmay be approximately 0.7. Advantageously, the hub to tip ratio isincreased relative to alternative configurations. Consequently, the coreinlet passage does not have to extend as far inwards for a given coreinlet radius, thereby allowing a shorter core inlet passage for a givenduct loading. Consequently, overall engine length and weight arereduced.

Each fan blade may be defined as having a radial span extending from aroot (or hub) at a radially inner gas-washed location, or 0% spanposition, to a tip at a 100% span position. The ratio of the radius ofthe fan blade at the hub to the radius of the fan blade at the tip maybe less than (or on the order of) any of: 0.4, 0.39, 0.38 0.37, 0.36,0.35, 0.34, 0.33, 0.32, 0.31, 0.3, 0.29, 0.28, 0.27, 0.26, or 0.25. Theratio of the radius of the fan blade at the hub to the radius of the fanblade at the tip may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). These ratios may commonly be referred to as the hub-to-tipratio. The radius at the hub and the radius at the tip may both bemeasured at the leading edge (or axially forwardmost) part of the blade.The hub-to-tip ratio refers, of course, to the gas-washed portion of thefan blade, i.e. the portion radially outside any platform.

The present disclosure is particularly applicable to engines having fanswith high hub to tip ratios, since this necessarily results in a highcore inlet radius. Consequently, in combination with a fan having a highhub to tip ratio, the present invention results in a significantlyshorter engine compared to conventional engines having high hub to tipratios, thereby resulting in significant weight saving and improvedpackaging.

The engine core may comprise an inter-compressor duct extending betweenan outlet of the low pressure compressor, and an inlet of the highpressure compressor.

The inter-compressor duct may define a second duct loading ΔR₂/L₂. Thesecond duct loading ΔR₂/L₂ may be defined by a second radius change ΔR₂divided by a second axial length L₂ of the inter-compressor duct. Thesecond radius change ΔR₂ may be defined by a difference between a radiusR_(IPC) of the inter-compressor passage measured at a mid-height of thetrailing edge of an axially rearmost low pressure compressor rotor at anaxially forward end of the inter-compressor passage, and a radiusR_(HPC) of the inter-compressor passage measured at a mid-height at aleading edge of a first rotor stage of the high pressure compressor. Thesecond duct loading may be between 0.3 and 0.6, may be between 0.35 and0.55, and may be approximately 0.5. Typically, the inter-compressor ducthas a higher duct loading than the core inlet passage. Consequently,although the larger radius final stage rotor of the low pressure rotorresults in a larger difference in radius between the forward and rearends of the inter-compressor duct, this can be accommodated withoutsignificantly increasing the length of the inter-compressor duct, inview of the higher duct loading that can be tolerated in this region.Consequently, overall engine length is reduced.

The engine may comprise a forward core mounting member which may extendbetween a radially inner core housing and a radially outer core housing.

The engine may comprise a plurality of outlet guide vanes locatedaxially rearward of the fan. The outlet guide vanes may be configured toprovide structural support for the engine core relative to a fanhousing.

The forward core mounting member may extend between an axial position ofthe inter-compressor duct and an axial position of a root trailing edgeof the outlet guide vanes. Advantageously, the relatively low number oflow pressure compressor stages permits a relatively short distancebetween the inter-compressor duct and fan outlet guide vanes (OGVs),which in turn permits a relatively short forward core mounting member,which further reduces weight. Furthermore, this permits the forward coremounting member to subtend a relatively small angle relative to theradial direction, thereby reducing bending stress in operation, and sofurther reducing weight and increasing strength.

The high pressure turbine may comprise two or fewer stages.

The low pressure turbine may comprise four or fewer stages and maycomprise three stages.

The gas turbine engine as described and/or claimed herein may have anysuitable general architecture. For example, the gas turbine engine mayhave further shafts that connect turbines and compressors, for examplethree shafts.

The gearbox is a reduction gearbox (in that the output to the fan is alower rotational rate than the input from the core shaft). Any type ofgearbox may be used. For example, the gearbox may be a “planetary” or“star” gearbox, as described in more detail elsewhere herein. Thegearbox may have any desired reduction ratio (defined as the rotationalspeed of the input shaft divided by the rotational speed of the outputshaft), for example greater than 2.5, for example in the range of from 3to 4.2, for example on the order of or at least 3, 3.1, 3.2, 3.3, 3.4,3.5, 3.6, 3.7, 3.8, 3.9, 4, 4.1 or 4.2. The gear ratio may be, forexample, between any two of the values in the previous sentence. Ahigher gear ratio may be more suited to “planetary” style gearbox. Insome arrangements, the gear ratio may be outside these ranges.

In any gas turbine engine as described and/or claimed herein, acombustor may be provided axially downstream of the fan and compressors.For example, the combustor may be directly downstream of (for example atthe exit of) the high pressure compressor. By way of further example,the flow at the exit to the combustor may be provided to the inlet ofthe high pressure turbine. The combustor may be provided upstream of theturbines.

Each compressor stage may comprise a row of rotor blades and a row ofstator vanes, which may be variable stator vanes (in that their angle ofincidence may be variable). The row of rotor blades and the row ofstator vanes may be axially offset from each other.

Similarly, each turbine may comprise any number of stages, for examplemultiple stages. Each stage may comprise a row of rotor blades and a rowof stator vanes. The row of rotor blades and the row of stator vanes maybe axially offset from each other.

The radius of the fan may be measured between the engine centreline andthe tip of a fan blade at its leading edge. The fan diameter (which maysimply be twice the radius of the fan) may be greater than (or on theorder of) any of: 250 cm (around 100 inches), 260 cm, 270 cm (around 105inches), 280 cm (around 110 inches), 290 cm (around 115 inches), 300 cm(around 120 inches), 310 cm, 320 cm (around 125 inches), 330 cm (around130 inches), 340 cm (around 135 inches), 350cm, 360cm (around 140inches), 370 cm (around 145 inches), 380 (around 150 inches) cm or 390cm (around 155 inches). The fan diameter may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds).

The rotational speed of the fan may vary in use. Generally, therotational speed is lower for fans with a higher diameter. Purely by wayof non-limitative example, the rotational speed of the fan at cruiseconditions may be less than 2500 rpm, for example less than 2300 rpm.Purely by way of further non-limitative example, the rotational speed ofthe fan at cruise conditions for an engine having a fan diameter in therange of from 250 cm to 300 cm (for example 250 cm to 280 cm) may be inthe range of from 1700 rpm to 2500 rpm, for example in the range of from1800 rpm to 2300 rpm, for example in the range of from 1900 rpm to 2100rpm. Purely by way of further non-limitative example, the rotationalspeed of the fan at cruise conditions for an engine having a fandiameter in the range of from 320 cm to 380 cm may be in the range offrom 1200 rpm to 2000 rpm, for example in the range of from 1300 rpm to1800 rpm, for example in the range of from 1400 rpm to 1600 rpm.

In use of the gas turbine engine, the fan (with associated fan blades)rotates about a rotational axis. This rotation results in the tip of thefan blade moving with a velocity U_(tip). The work done by the fanblades 13 on the flow results in an enthalpy rise dH of the flow. A fantip loading may be defined as dH/U_(tip) ², where dH is the enthalpyrise (for example the 1-D average enthalpy rise) across the fan andU_(tip) is the (translational) velocity of the fan tip, for example atthe leading edge of the tip (which may be defined as fan tip radius atleading edge multiplied by angular speed). The fan tip loading at cruiseconditions may be greater than (or on the order of) any of: 0.3, 0.31,0.32, 0.33, 0.34, 0.35, 0.36, 0.37, 0.38, 0.39 or 0.4 (all units in thisparagraph being Jkg⁻¹ K⁻¹/(ms⁻¹)²). The fan tip loading may be in aninclusive range bounded by any two of the values in the previoussentence (i.e. the values may form upper or lower bounds).

Gas turbine engines in accordance with the present disclosure may haveany desired bypass ratio, where the bypass ratio is defined as the ratioof the mass flow rate of the flow through the bypass duct to the massflow rate of the flow through the core at cruise conditions. In somearrangements the bypass ratio may be greater than (or on the order of)any of the following: 10, 10.5, 11, 11.5, 12, 12.5, 13, 13.5, 14, 14.5,15, 15.5, 16, 16.5, 17, 17.5, 18, 18.5, 19, 19.5 or 20. The bypass ratiomay be in an inclusive range bounded by any two of the values in theprevious sentence (i.e. the values may form upper or lower bounds). Thebypass duct may be substantially annular. The bypass duct may beradially outside the core engine. The radially outer surface of thebypass duct may be defined by a nacelle and/or a fan case.

The overall core pressure ratio of a gas turbine engine as describedand/or claimed herein may be defined as the ratio of the stagnationpressure upstream of the low pressure compressor to the stagnationpressure at the exit of the highest pressure compressor (before entryinto the combustor).

Specific thrust of an engine may be defined as the net thrust of theengine divided by the total mass flow through the engine. At cruiseconditions, the specific thrust of an engine described and/or claimedherein may be less than (or on the order of) any of the following: 110Nkg⁻¹ s, 105 Nkg⁻¹ s, 100 Nkg⁻¹ s, 95 Nkg⁻¹ s, 90 Nkg⁻¹ s, 85 Nkg⁻¹ s or80 Nkg⁻¹ s. The specific thrust may be in an inclusive range bounded byany two of the values in the previous sentence (i.e. the values may formupper or lower bounds). Such engines may be particularly efficient incomparison with conventional gas turbine engines.

A gas turbine engine as described and/or claimed herein may have anydesired maximum thrust. Purely by way of non-limitative example, a gasturbine as described and/or claimed herein may be capable of producing amaximum thrust of at least (or on the order of) any of the following:160 kN, 170 kN, 180 kN, 190 kN, 200 kN, 250 kN, 300 kN, 350 kN, 400 kN,450 kN, 500 kN, or 550 kN. The maximum thrust may be in an inclusiverange bounded by any two of the values in the previous sentence (i.e.the values may form upper or lower bounds). The thrust referred to abovemay be the maximum net thrust at standard atmospheric conditions at sealevel plus 15 deg C. (ambient pressure 101.3 kPa, temperature 30 degC.), with the engine static.

In use, the temperature of the flow at the entry to the high pressureturbine may be particularly high. This temperature, which may bereferred to as TET, may be measured at the exit to the combustor, forexample immediately upstream of the first turbine vane, which itself maybe referred to as a nozzle guide vane. At cruise, the TET may be atleast (or on the order of) any of the following: 1400K, 1450K, 1500K,1550K, 1600K or 1650K. The TET at cruise may be in an inclusive rangebounded by any two of the values in the previous sentence (i.e. thevalues may form upper or lower bounds). The maximum TET in use of theengine may be, for example, at least (or on the order of) any of thefollowing: 1700K, 1750K, 1800K, 1850K, 1900K, 1950K or 2000K. Themaximum TET may be in an inclusive range bounded by any two of thevalues in the previous sentence (i.e. the values may form upper or lowerbounds). The maximum TET may occur, for example, at a high thrustcondition, for example at a maximum take-off (MTO) condition.

A fan blade and/or aerofoil portion of a fan blade described and/orclaimed herein may be manufactured from any suitable material orcombination of materials. For example at least a part of the fan bladeand/or aerofoil may be manufactured at least in part from a composite,for example a metal matrix composite and/or an organic matrix composite,such as carbon fibre. By way of further example at least a part of thefan blade and/or aerofoil may be manufactured at least in part from ametal, such as a titanium based metal or an aluminium based material(such as an aluminium-lithium alloy) or a steel based material. The fanblade may comprise at least two regions manufactured using differentmaterials. For example, the fan blade may have a protective leadingedge, which may be manufactured using a material that is better able toresist impact (for example from birds, ice or other material) than therest of the blade. Such a leading edge may, for example, be manufacturedusing titanium or a titanium-based alloy. Thus, purely by way ofexample, the fan blade may have a carbon-fibre or aluminium based body(such as an aluminium lithium alloy) with a titanium leading edge.

A fan as described and/or claimed herein may comprise a central portion,from which the fan blades may extend, for example in a radial direction.The fan blades may be attached to the central portion in any desiredmanner. For example, each fan blade may comprise a fixture which mayengage a corresponding slot in the hub (or disc). Purely by way ofexample, such a fixture may be in the form of a dovetail that may slotinto and/or engage a corresponding slot in the hub/disc in order to fixthe fan blade to the hub/disc. By way of further example, the fan bladesmaybe formed integrally with a central portion. Such an arrangement maybe referred to as a bladed disc or a bladed ring. Any suitable methodmay be used to manufacture such a bladed disc or bladed ring. Forexample, at least a part of the fan blades may be machined from a blockand/or at least part of the fan blades may be attached to the hub/discby welding, such as linear friction welding.

The gas turbine engines described and/or claimed herein may or may notbe provided with a variable area nozzle (VAN). Such a variable areanozzle may allow the exit area of the bypass duct to be varied in use.The general principles of the present disclosure may apply to engineswith or without a VAN.

The fan of a gas turbine as described and/or claimed herein may have anydesired number of fan blades, for example 14, 16, 18, 20, 22, 24 or 26fan blades.

As used herein, cruise conditions have the conventional meaning andwould be readily understood by the skilled person. Thus, for a given gasturbine engine for an aircraft, the skilled person would immediatelyrecognise cruise conditions to mean the operating point of the engine atmid-cruise of a given mission (which may be referred to in the industryas the “economic mission”) of an aircraft to which the gas turbineengine is designed to be attached. In this regard, mid-cruise is thepoint in an aircraft flight cycle at which 50% of the total fuel that isburned between top of climb and start of descent has been burned (whichmay be approximated by the midpoint—in terms of time and/ordistance—between top of climb and start of descent. Cruise conditionsthus define an operating point of, the gas turbine engine that providesa thrust that would ensure steady state operation (i.e. maintaining aconstant altitude and constant Mach Number) at mid-cruise of an aircraftto which it is designed to be attached, taking into account the numberof engines provided to that aircraft. For example where an engine isdesigned to be attached to an aircraft that has two engines of the sametype, at cruise conditions the engine provides half of the total thrustthat would be required for steady state operation of that aircraft atmid-cruise.

In other words, for a given gas turbine engine for an aircraft, cruiseconditions are defined as the operating point of the engine thatprovides a specified thrust (required to provide—in combination with anyother engines on the aircraft—steady state operation of the aircraft towhich it is designed to be attached at a given mid-cruise Mach Number)at the mid-cruise atmospheric conditions (defined by the InternationalStandard Atmosphere according to ISO 2533 at the mid-cruise altitude).For any given gas turbine engine for an aircraft, the mid-cruise thrust,atmospheric conditions and Mach Number are known, and thus the operatingpoint of the engine at cruise conditions is clearly defined.

Purely by way of example, the forward speed at the cruise condition maybe any point in the range of from Mach 0.7 to 0.9, for example 0.75 to0.85, for example 0.76 to 0.84, for example 0.77 to 0.83, for example0.78 to 0.82, for example 0.79 to 0.81, for example on the order of Mach0.8, on the order of Mach 0.85 or in the range of from 0.8 to 0.85. Anysingle speed within these ranges may be part of the cruise condition.For some aircraft, the cruise conditions may be outside these ranges,for example below Mach 0.7 or above Mach 0.9.

Purely by way of example, the cruise conditions may correspond tostandard atmospheric conditions (according to the International StandardAtmosphere, ISA) at an altitude that is in the range of from 10000 m to15000 m, for example in the range of from 10000 m to 12000 m, forexample in the range of from 10400 m to 11600 m (around 38000 ft), forexample in the range of from 10500 m to 11500 m, for example in therange of from 10600 m to 11400 m, for example in the range of from 10700m (around 35000 ft) to 11300 m, for example in the range of from 10800 mto 11200 m, for example in the range of from 10900 m to 11100 m, forexample on the order of 11000 m. The cruise conditions may correspond tostandard atmospheric conditions at any given altitude in these ranges.

Purely by way of example, the cruise conditions may correspond to anoperating point of the engine that provides a known required thrustlevel (for example a value in the range of from 30 kN to 35 kN) at aforward Mach number of 0.8 and standard atmospheric conditions(according to the International Standard Atmosphere) at an altitude of38000 ft (11582 m). Purely by way of further example, the cruiseconditions may correspond to an operating point of the engine thatprovides a known required thrust level (for example a value in the rangeof from 50 kN to 65 kN) at a forward Mach number of 0.85 and standardatmospheric conditions (according to the International StandardAtmosphere) at an altitude of 35000 ft (10668 m).

In use, a gas turbine engine described and/or claimed herein may operateat the cruise conditions defined elsewhere herein. Such cruiseconditions may be determined by the cruise conditions (for example themid-cruise conditions) of an aircraft to which at least one (for example2 or 4) gas turbine engine may be mounted in order to provide propulsivethrust.

According to an aspect, there is provided a method of operating a gasturbine engine as described and/or claimed herein. The operation may beat the cruise conditions as defined elsewhere herein (for example interms of the thrust, atmospheric conditions and Mach Number).

According to an aspect, there is provided a method of operating anaircraft comprising a gas turbine engine as described and/or claimedherein. The operation according to this aspect may include (or may be)operation at the mid-cruise of the aircraft, as defined elsewhereherein.

The method may comprise, at cruise conditions, operating the low andhigh pressure compressors to provide a pressure ratio between 30:1 and50:1.

The skilled person will appreciate that except where mutually exclusive,a feature or parameter described in relation to any one of the aboveaspects may be applied to any other aspect. Furthermore, except wheremutually exclusive, any feature or parameter described herein may beapplied to any aspect and/or combined with any other feature orparameter described herein.

Embodiments will now be described by way of example only, with referenceto the Figures, in which:

FIG. 1 is a sectional side view of a gas turbine engine in accordancewith the present disclosure;

FIG. 2 is a close up sectional side view of an upstream portion of thegas turbine engine of FIG. 1;

FIG. 3 is a partially cut-away view of a gearbox for a gas turbineengine; and

FIG. 4 is a sectional side view of a gas turbine engine not inaccordance with the present disclosure.

FIG. 1 illustrates a gas turbine engine 10 having a principal rotationalaxis 9. The engine 10 comprises an air intake 12 and a propulsive fan 23that generates two airflows: a core airflow A and a bypass airflow B.The gas turbine engine 10 comprises a core 11 that receives the coreairflow A. The engine core 11 comprises, in axial flow series, a lowpressure compressor 14, a high-pressure compressor 15, combustionequipment 16, a high-pressure turbine 17, a low pressure turbine 19 anda core exhaust nozzle 20. A nacelle 21 surrounds the gas turbine engine10 and defines a bypass duct 22 and a bypass exhaust nozzle 18. Thebypass airflow B flows through the bypass duct 22. The fan 23 isattached to and driven by the low pressure turbine 19 via a shaft 26 andan epicyclic gearbox 30. A plurality of Outlet Guide Vanes (OGV) 56 areprovided downstream of the fan 23, within the nacelle 21.

In use, the core airflow A is accelerated and compressed by the lowpressure compressor 14 and directed into the high pressure compressor 15where further compression takes place. The compressed air exhausted fromthe high pressure compressor 15 is directed into the combustionequipment 16 where it is mixed with fuel and the mixture is combusted.The resultant hot combustion products then expand through, and therebydrive, the high pressure and low pressure turbines 17, 19 before beingexhausted through the nozzle 20 to provide some propulsive thrust. Thehigh pressure turbine 17 drives the high pressure compressor 15 by asuitable interconnecting shaft 27. The fan 23 generally provides themajority of the propulsive thrust. The epicyclic gearbox 30 is areduction gearbox.

A front part of the engine 10 is shown in more detail. The low pressureturbine 19 drives the shaft 26, which is coupled to a sun wheel, or sungear, 28 of the epicyclic gear arrangement 30, which is also shown inmore detail in FIG. 3. Radially outwardly of the sun gear 28 andintermeshing therewith is a plurality of planet gears 32 that arecoupled together by a planet carrier 34. The planet carrier 34constrains the planet gears 32 to precess around the sun gear 28 insynchronicity whilst enabling each planet gear 32 to rotate about itsown axis. The planet carrier 34 is coupled via linkages 36 to the fan 23in order to drive its rotation about the engine axis 9. Radiallyoutwardly of the planet gears 32 and intermeshing therewith is anannulus or ring gear 38 that is coupled, via linkages 40, to astationary supporting structure 24.

Note that the terms “low pressure turbine” and “low pressure compressor”as used herein may be taken to mean the lowest pressure turbine stagesand lowest pressure compressor stages (i.e. not including the fan 23)respectively and/or the turbine and compressor stages that are connectedtogether by the interconnecting shaft 26 with the lowest rotationalspeed in the engine (i.e. not including the gearbox output shaft thatdrives the fan 23). In some literature, the “low pressure turbine” and“low pressure compressor” referred to herein may alternatively be knownas the “intermediate pressure turbine” and “intermediate pressurecompressor”. Where such alternative nomenclature is used, the fan 23 maybe referred to as a first, or lowest pressure, compression stage.

The epicyclic gearbox 30 is shown by way of example in greater detail inFIG. 3. Each of the sun gear 28, planet gears 32 and ring gear 38comprise teeth about their periphery to intermesh with the other gears.However, for clarity only exemplary portions of the teeth areillustrated in FIG. 3. There are four planet gears 32 illustrated,although it will be apparent to the skilled reader that more or fewerplanet gears 32 may be provided within the scope of the claimedinvention. Practical applications of a planetary epicyclic gearbox 30generally comprise at least three planet gears 32.

The epicyclic gearbox 30 illustrated by way of example in FIGS. 2 and 3is of the planetary type, in that the planet carrier 34 is coupled to anoutput shaft via linkages 36, with the ring gear 38 fixed. However, anyother suitable type of epicyclic gearbox 30 may be used. By way offurther example, the epicyclic gearbox 30 may be a star arrangement, inwhich the planet carrier 34 is held fixed, with the ring (or annulus)gear 38 allowed to rotate. In such an arrangement the fan 23 is drivenby the ring gear 38. By way of further alternative example, the gearbox30 may be a differential gearbox in which the ring gear 38 and theplanet carrier 34 are both allowed to rotate.

It will be appreciated that the arrangement shown in FIGS. 2 and 3 is byway of example only, and various alternatives are within the scope ofthe present disclosure. Purely by way of example, any suitablearrangement may be used for locating the gearbox 30 in the engine 10and/or for connecting the gearbox 30 to the engine 10. By way of furtherexample, the connections (such as the linkages 36, 40 in the FIG. 2example) between the gearbox 30 and other parts of the engine 10 (suchas the input shaft 26, the output shaft and the fixed structure 24) mayhave any desired degree of stiffness or flexibility. By way of furtherexample, any suitable arrangement of the bearings between rotating andstationary parts of the engine (for example between the input and outputshafts from the gearbox and the fixed structures, such as the gearboxcasing) may be used, and the disclosure is not limited to the exemplaryarrangement of FIGS. 1 to 3. For example, where the gearbox 30 has astar arrangement (described above), the skilled person would readilyunderstand that the arrangement of output and support linkages andbearing locations would typically be different to that shown by way ofexample in FIG. 2.

Accordingly, the present disclosure extends to a gas turbine enginehaving any arrangement of gearbox styles (for example star orplanetary), support structures, input and output shaft arrangement, andbearing locations.

Optionally, the gearbox may drive additional and/or alternativecomponents (e.g. the intermediate pressure compressor and/or a boostercompressor).

Referring once more to FIGS. 1 and 2, each of the compressors 14, 15comprises a multi-stage axial flow compressor.

Referring now to FIG. 2, the low pressure compressor 14 consists ofexactly two stages 41, 42—i.e. no more than or less than two stages.Each stage 41, 42 comprises at least one respective compressor rotor 43a, 43 b, and may comprise a respective stator 44 a, 44 b. The respectiverotor 43 a, 43 b and stator 44 a, 44 b are generally axially spaced. Inthe present case, the first stator 44 a is upstream of the first rotor43 a. One or more further stators such as an inlet stator 45 may beprovided—however, since no additional rotor is associated with the inletstator 45, this does not constitute an additional stage, since nopressure rise is provided by the inlet stator 45 alone. As will beappreciated by the person skilled in the art, the rotors 43 a, 43 b arecoupled to the respective shaft (i.e. the low pressure shaft 26 in thecase of the low pressure compressor 15) by corresponding discs 46, andso turn with the shaft 26. On the other hand, the stators 44 a, 44 b areheld stationary. In some cases, the stators 44 a, 44 b may pivot abouttheir long axes, to adjust the angle of attack and inlet and outlet areafor the respective compressor stage. Such stators are known as “variablestator vanes” or VSVs.

The high pressure compressor 15 similarly comprises either nine or tenstages, and in the described embodiment consists of ten stages. Thefirst two stages 47, 48 are shown in FIG. 1, while the remainder are notshown to aid clarity.

Between them, the high and low pressure compressors 15, 16 define acruise in use overall core pressure ratio (OPR). The core OPR is definedas the ratio of the stagnation pressure upstream of the first stage 44of the low pressure compressor 15 to the stagnation pressure at the exitof the highest pressure compressor 16 (before entry into the combustor).The core OPR excludes any pressure rise generated by the fan 23 wherethe fan provides air flow to the core, so a total engine overallpressure ratio (EPR) may be higher than the core OPR. In the presentdisclosure, the overall core OPR is between 30 and 50 at cruiseconditions, as defined above. In the described embodiment, the core OPRis 40, and may take any value between these upper and low bounds. Forexample, the core OPR may be any of 35, 40, 45, and 50 at cruiseconditions.

As will be understood, the core OPR will vary according to atmospheric,flight and engine conditions. However, the cruise OPR (i.e. the highestachievable OPR for that engine) will occur at a particular point in theflight cycle for a given engine design.

As will be understood, a large design space must be considered whendesigning a gas turbine engine to determine an optimal engine withrespect to a chosen metric (such as engine weight, cost, thermalefficiency, propulsive efficiency, or a balance of these). In manycases, there may be a large number of feasible solutions for a given setof conditions to achieve a desired metric.

One such variable is core OPR. As core OPR increases, thermal efficiencyalso tends to increase, and so a high OPR is desirable. Even once aparticular OPR is chosen however, a number of design variables must bechosen to meet the chosen OPR.

In choosing a core OPR, a further design variable is the amount ofpressure rise provided by the low pressure compressor 15 relative tothat provided by the high pressure compressor 16 (sometimes referred toas “worksplit”). As will be understood, the total core OPR can bedetermined by multiplying the low pressure compressor pressure ratio(i.e. the ratio between the stagnation pressure at the outlet of the lowpressure compressor to the stagnation pressure at the inlet of the lowpressure compressor 15) by the high pressure compressor ratio (i.e. theratio between the stagnation pressure at the outlet of the high pressurecompressor 16 to the stagnation pressure at the inlet of the highpressure compressor 16). Consequently, a higher core OPR can be providedby increasing the high pressure compressor ratio, the low pressurecompressor ratio, or both.

The inventors have found that a particularly efficient work split for agas turbine engine having a core OPR in the above described range can beprovided by providing a high pressure compressor 16 having a pressureratio of between 17:1 and 25:1. In the present example, the highpressure compressor has a pressure ratio of approximately 20:1. It hasbeen found to be difficult to provide a pressure ratio significantlygreater than 25:1 on a compressor provided on a single shaft usingcurrent technology. Consequently, to provide the necessary core OPR, alow pressure compressor ratio of between 1.5:1 and 2:1 is required. Inthe present example, the low pressure compressor 15 has a pressure ratioof approximately 1.7:1, giving a core OPR of 34:1.

Similarly, there are a number of ways to increase the compressorpressure ratio. A first method is to increase the stage loading. Stageloading is defined as the stagnation pressure ratio across an individualstage (rotor and stator) of a compressor. Similarly, an average stageloading can be defined as the sum of the stage loadings of eachcompressor stage of a compressor, divided by the number of stages. Forexample, in the present disclosure, the average stage loading of the lowpressure compressor 15 is 1.3. This can in turn be increased by one ormore of increasing the rotor speed at the cruise compression conditions,increasing the turning provided by the blades, or increasing the radiusof the tips of the compressor rotors, which in turn necessitates anincrease in the radius of the roots of the compressor rotors to maintaina given flow area. Each of these options has associated advantages anddisadvantages. For instance, increasing low pressure compressor rotorspeed necessitates either an increase in the reduction ratio of thegearbox 30, or a reduction in the fan 23 radius, in order to maintainfan tip speeds at a desired level for noise and efficiency reasons. Onthe other hand, increasing the compressor tip radius necessitates anincrease in weight, in view of the larger compressor discs that arerequired. Increased turning of the airflow may result in lower surgemargin, and reduced efficiency. In any case, a higher stage loading mayresult in a lower efficiency, since the increased rotor tip speed orhigher turning leads to lower compressor efficiencies, in view of lossesassociated with aerodynamic shocks as the tips significantly exceed thespeed of sound.

A second option is to increase the number of stages in the respectivecompressors, thereby maintaining a low stage loading, low rotationalspeed, and low disc weight. Again, this can be achieved by adding astage to either the low pressure compressor 15 or high pressurecompressor 16. However, this will generally result in a higher weightand cost associated with the additional stage.

A further complication is the presence of the gearbox 30. The gearboxprovides additional design freedom, since, as noted above, the gearboxreduction ratio can be selected to provide a preferred fan tip speedindependently of both fan radius and low pressure compressor rotorspeed. However, the gearbox also presents constraints in view of itslarge size. Consequently, the large radius required radially inward ofthe fan 23 inherent in a geared turbofan having an epicyclic gearboxdictates a fan 23 having a large hub radius, i.e. a large radialdistance between the engine centre 9 and the aerodynamic root of the fanblades 23. Furthermore, in view of the relatively slow turning fantypical of geared turbofans, relatively little pressure rise is providedby the inner radius of the fan 23, and so geared turbofans tend to havea high hub to tip ratio.

As can be seen from FIG. 2, a core inlet passage 49 is provided. Thecore inlet passage 49 is defined by radially inner 50 and outer 51walls, which enclose the core flow A. The core inlet passage 49 definesa core inlet 52 at an axial position of a leading edge 52 of the enginesection stator 45, and terminates at a leading edge of the first rotorstage 44 of the low pressure compressor.

Similarly, an inter-compressor duct 53 is provided, which is defined bythe radially inner 50 and outer 51 walls. The inter-compressor duct 53extends between an outlet of the low pressure compressor 15 and an inletof the high pressure compressor 16.

As can be seen, there is a mismatch between a radius R_(inlet) of thecore inlet passage at the inlet relative to the radius R_(outlet) at thetermination of the core inlet passage 49. The radius of the inletR_(inlet) can be determined by measuring the radial distance between amid-span position (i.e. equidistance between a root and a tip of theaerofoil portion) of the leading edge of the engine section stator 45,and the rotational axis 9 of the engine. Similarly, the R_(outlet) canbe determined by measuring the radial distance between a mid-spanposition (i.e. equidistance between a root and a tip of the aerofoilportion) of the leading edge of the first compressor rotor 44 a of thelow pressure compressor 15, and the rotational axis 9 of the engine, asshown in FIG. 2. The difference between these radii ΔR₁ divided by anaxial length L₁ of the inlet passage 49 defines a first “duct loading”ΔR₁/L₁. For given inlet and outlet flow conditions, a maximum ductloading is required, otherwise flow separation will occur at theradially inner wall 50, resulting in turbulence at the first rotor stage44, and low engine performance. Consequently the radius difference ΔR₁has been found to significantly affect engine length. Engine length inturn is a significant driver of engine weight, since any increase inengine length results in a significant increase in structure, as well asincreased shaft lengths. Increased engine length may have other negativeconsequences, such as difficulty in installing the engine within spaceconstraints on the aircraft, and may even have impacts on other areas,such as wing flutter, due to the change in centre of mass. Consequently,it is desirable to minimise engine length.

The inventors have determined that a particular combination ofcompressor parameters can result in reduced engine length, whileproviding a highly efficient, high pressure ratio core.

As noted previously, a low pressure compressor 15 having two compressorstages is chosen. In principle, any number of low and pressurecompressor 15 stages could be chosen to provide the required OPR.However, the inventors have found that providing a low pressurecompressor having two stages provides distinct advantages.

To achieve the required pressure rise over only two stages, a relativelylarge diameter first stage 41 is chosen, with relatively little airflowturning per stage. In particular, a first stage rotor 43 having a hub totip ratio of between 0.5 and 0.7 is chosen, and in the presentembodiment, the hub to tip ratio is approximately 0.65. As noted above,this may result in high stage weight and relatively low stage efficiencyin view of the high tip speed—however, this has been found to be morethan compensated for by the reduced engine length.

As can be seen, the large first stage rotor 44 radius results in arelatively low difference between the inlet radius R_(inlet) and theoutlet radius, R_(outlet), and so a relatively short inlet passage 49for a given duct loading. The inventors have found that, for a gearedturbofan having a relatively slow turning fan, which develops relativelylittle pressure at the core inlet, a first duct loading of between 0.3and 0.6 can be tolerated. A first duct loading between 0.35 and 0.55 isfound to provide adequate resistance to flow separate over a wide rangeof conditions, and may be chosen to decrease the risk of enginecompressor surge. For example, in the present embodiment, the first ductloading is approximately 0.5

A relatively high stage count (nine or ten stages) is chosen for thehigh pressure compressor 15, and a relatively low radius high pressurecompressor 15 first stage rotor is chosen. This can be selected in viewof the relatively high rotational speed of the high pressure shaft 27.

This high speed, low radius first stage of the high pressure compressor15, in combination with the relatively large radius of the second stage42 of the low pressure compressor 14, results in a relatively largesecond duct loading for the inter-compressor duct 53. This in turnresults in a longer inter-compressor duct compared to a three stage lowpressure compressor, which partly offsets the advantages of thisarrangement. However, in view of the higher pressure, higher velocityair at this point, a higher duct loading can be achieved, and so theincrease in overall engine length due to this effect is modest.

As will be understood, the second duct loading can be defined by asecond radius change ΔR₂ divided by a second axial length L₂ of theinter-compressor duct. The second radius change ΔR₂ can be defined by adifference between a radius R_(IPC) of the inter-compressor passage 53measured at a mid-height of the trailing edge the axially rearmost lowpressure compressor rotor 43 b at an axially forward end of theinter-compressor passage 53, and an outer radius R_(HPC) of theinter-compressor passage measured at a mid-height at a leading edge of afirst rotor stage of the high pressure compressor 16. The second ductloading is generally higher than the first duct loading. This may bebetween 0.3 and 0.6, may be between 0.35 and 0.55, and may beapproximately 0.5.

To drive the ten stage high pressure compressor 15, a two stage highpressure turbine 17 may be necessary. Again, the number of turbinestages can be determined in a similar manner to the number of compressorstages. Similarly, to drive the low pressure compressor 14 and fan 23, athree or four stage low pressure turbine 19 is provided.

Further advantages are achieved in view of the reduced core inletpassage 49. As can be seen in FIG. 1, a forward core mounting 54 isprovided. The forward core mounting 54 provides support for the corerelative to the nacelle 21. The forward core mounting 54 extends betweenthe core outer wall 51 at an axial position corresponding to theinter-compressor duct 53, to a core outer housing 55 at an axialposition of the OGV 56, close to a trailing edge 57 of the OGV 56.

In view of the relatively short low pressure compressor 14, and shortcore inlet 49, the axial distance between the OGV 56 andinter-compressor duct 53 is minimised, thereby reducing the weight ofthe core mounting 54. Furthermore, the angle relative to the radialplane is reduced, thereby reducing the bending moment applied due toradial forces. Consequently, the weight of the core mounting 54 can bereduced further.

The advantages of the disclosed engine can be seen when comparing theengine of FIG. 1 to an alternative engine configuration 110 shown inFIG. 4.

The engine shown in FIG. 4 is similar to that shown in FIGS. 1 to 3, andonly differences will be described. However, the engine 110 is notwithin the scope of the present disclosure, since a low pressurecompressor 114 is provided having three compressor stages.

The overall pressure ratio, work split, gearbox ratio and shaftrotational speed of the engine 110 is however maintained at the samevalues as the engine 10. Consequently, in this example, the hub to tipratio of the low pressure compressor 114 is altered, with the lowpressure compressor 114 having a reduced diameter. Consequently, a coreinlet 149 is lengthened, in view of the high duct loading. As a result,both the low pressure compressor length and the core inlet duct lengthis increased, resulting in a large increase in overall engine length.Conceptual design and modelling has been undertaken, which has revealedthat such a design would be expected to have a length which isapproximately 10% longer and heavier than the engine 10 of FIG. 1.Similarly, both the length and angle relative to the radial direction ofthe core engine mount 154 is significantly increased, thereby leading tofurther weight increases.

It will be understood that the invention is not limited to theembodiments above-described and various modifications and improvementscan be made without departing from the concepts described herein. Exceptwhere mutually exclusive, any of the features may be employed separatelyor in combination with any other features and the disclosure extends toand includes all combinations and sub-combinations of one or morefeatures described herein.

Other gas turbine engines to which the present disclosure may be appliedmay have alternative configurations. For example, such engines may havean alternative number of compressors and/or turbines and/or analternative number of interconnecting shafts. By way of further example,the gas turbine engine shown in FIG. 1 has a split flow nozzle 18, 20meaning that the flow through the bypass duct 22 has its own nozzle 18that is separate to and radially outside the core engine nozzle 20.However, this is not limiting, and any aspect of the present disclosuremay also apply to engines in which the flow through the bypass duct 22and the flow through the core 11 are mixed, or combined, before (orupstream of) a single nozzle, which may be referred to as a mixed flownozzle. One or both nozzles (whether mixed or split flow) may have afixed or variable area. Whilst the described example relates to aturbofan engine, the disclosure may apply, for example, to any type ofgas turbine engine, such as an open rotor (in which the fan stage is notsurrounded by a nacelle) or turboprop engine, for example.

The geometry of the gas turbine engine 10, and components thereof, isdefined by a conventional axis system, comprising an axial direction(which is aligned with the rotational axis 9), a radial direction (inthe bottom-to-top direction in FIG. 1), and a circumferential direction(perpendicular to the page in the FIG. 1 view). The axial, radial andcircumferential directions are mutually perpendicular.

1. A gas turbine engine comprising a gas turbine engine core comprisinghigh and low pressure compressors and high and low pressure turbines,the high pressure compressor and high pressure turbine being coupled bya high pressure shaft, and the low pressure compressor and low pressureturbine being coupled by a low pressure shaft, the engine furthercomprising a fan coupled to the low pressure shaft by a reductiongearbox, wherein the low pressure compressor comprises no more than twocompressor stages, and the low and high pressure compressor togetherdefine a cruise overall core pressure ratio of between 30 and
 50. 2. Agas turbine engine according to claim 1, wherein the high pressurecompressor defines a cruise overall pressure ratio of between 16:1 and27:1.
 3. A gas turbine engine according to claim 1, wherein the highpressure compressor has no fewer than 8 stages and no more than 12stages.
 4. A gas turbine engine according to claim 1, wherein the lowpressure compressor comprises a cruise pressure ratio of between 1.5 and2.
 5. A gas turbine engine according to claim 1, wherein the enginecomprises a core inlet passage extending between a core inlet and aninlet of the low pressure compressor, a first radius change ΔR₁ beingdefined by a difference between a radius (R_(inlet)) at a mid-height ofthe leading edge of an engine section stator aerofoil at an axiallyforward end of the core inlet passage, and an outlet radius R_(outlet)of the core inlet passage measured at a mid-height at a leading edge ofa first rotor stage of the low pressure compressor, a first duct loadingΔR₁/L₁ being defined by a ratio of the first radius change ΔR₁ to afirst axial length L of the inlet duct between the axially forward endand the leading edge of the first stage of the low pressure compressor,wherein the first duct loading is between 0.3 and 0.6.
 6. A gas turbineengine according to claim 1, wherein the first compressor rotor of thelow pressure compressor defines a hub to tip ratio defined by a radiusof a tip of a leading edge of the compressor rotor divided by a radiusof a root at the leading edge of the compressor rotor, wherein the hubto tip ratio is between 0.6 and 0.75.
 7. A gas turbine engine accordingto claim 1, wherein the engine core comprises an inter-compressor ductextending between an outlet of the low pressure compressor and an inletof the high pressure compressor, the inter-compressor duct defining asecond duct loading ΔR₂/L₂ comprising a second radius change ΔR₂ dividedby a second axial length L₂ of the inter-compressor duct, wherein thesecond radius change ΔR₂ can be determined by a difference between aradius R_(IPC) of the inter-compressor passage measured at a mid-heightof the trailing edge of an axially rearmost low pressure compressorrotor at an axially forward end of the inter-compressor passage, and aradius R_(HPC) of the inter-compressor passage measured at a mid-heightat a leading edge of a first rotor stage of the high pressurecompressor, wherein the second duct loading is between 0.3 and 0.6.
 8. Agas turbine engine according to claim 1, wherein the engine comprises aforward core mounting member which extends between a radially inner corehousing and a radially outer core housing.
 9. A gas turbine engineaccording to claim 1, wherein the engine comprises a plurality of outletguide vanes located axially rearward of the fan, wherein the outletguide vanes are configured to provide structural support for the enginecore relative to a fan housing.
 10. A gas turbine engine according toclaim 8, wherein the forward core mounting member extends between anaxial position of the inter-compressor duct and an axial position of aroot trailing edge of the outlet guide vanes.
 11. A method of operatinga gas turbine engine in accordance with claim 1, comprising, at cruiseconditions, operating the low and high pressure compressors to provide apressure ratio between 30:1 and 50:1.
 12. A gas turbine engine accordingto claim 9, wherein the forward core mounting member extends between anaxial position of the inter-compressor duct and an axial position of aroot trailing edge of the outlet guide vanes.
 13. A gas turbine engineaccording to claim 2, wherein the high pressure compressor defines acruise overall pressure ratio of between 17:1 and 20:1.
 14. A gasturbine engine according to claim 3, wherein the high pressurecompressor comprises 9 or 10 stages.
 15. A gas turbine engine accordingto claim 5, wherein the first duct loading is between 0.35 and 0.55. 16.A gas turbine engine according to claim 5, wherein the first ductloading is approximately 0.4.
 17. A gas turbine engine according toclaim 6, wherein the hub to tip ratio is approximately 0.7.
 18. A gasturbine engine according to claim 7, wherein the second duct loading isbetween 0.35 and 0.55.
 19. A gas turbine engine according to claim 7,wherein the second duct loading is approximately 0.5.